A reversible flight control system is where there is a direct mechanical linkage connection between the control lever in the flight compartment and the flight control surface. In a reversible system, if the surface is moved the control lever will move. In reversible systems, the pilot directly feels hinge moments applied to the surface. A simple example of a reversible flight control system is shown in Figure 1.
Figure 1 Simple Reversible Flight Control System
A simple schematic of a reversible flight control system is shown in Figure 2. Schematically, the system shown in Figure 2 is identical to the system shown in Figure 1. The representation shown in Figure 2 is used for analysis. In Figure 2, the control stick, pushrod and O2 sector are a four bar linkage. The O3 sector, pushrod and O4 bellcrank are a four bar linkage. The two four bar linkages are connected using a cable system. Analysis of the system would require an analysis of both four bar linkage arrangements plus the cable system. The cable system acts like springs that connect the two sectors.
Figure 2 Analytical Representation of System Shown in Figure 1
Reversible flight control systems are used on smaller aircraft where the hinge moment (surface) loads are small enough that a mechanical linkage system is adequate. For a reversible system, the maximum force input is limited by the capability of a pilot and the range of control input. For example, rudder pedal travel is limited to 5 inches in each direction. If the rudder must move ±30 degrees, then the gear ratio is 5/30 = 0.167 inch / degree = 9.55 inch / radian = 9.55 in-lbs / lb (see Mechanism – Mechanical Advantage). Using FAA guidelines per 14CFR 25.143(d), the maximum short-term rudder pedal force is limited to 150 lbs. For this example, the maximum rudder hinge moment is then (150 lbs) x (9.55 in-lbs / lb) = 1432.5 in-lbs. If the required hinge moment is greater than 1432.5 in-lbs, some type of power assist or irreversible system is required.
In some reversible systems, high surface hinge moments are only required under certain conditions – usually a failure condition – and some type of boost mechanism is implemented for this certain condition. This saves the cost and weight of going with an irreversible system. A good example is the increase in rudder control force required to overcome an engine out scenario. An engine out failure creates a yaw moment on the aircraft that must be offset with rudder deflection. For example, the maximum normal operating forces for a rudder system may be 600 in-lbs (83 lbs rudder pedal force), but for an engine out situation the rudder hinge moment might increase to 1200 in-lbs (166 lbs rudder pedal force). Since the 1200 in-lbs exceeds the maximum allowable force of 150 lbs rudder pedal force, the basic design would not be acceptable. Some methods to overcome this situation without implementing a fully powered system include (1) rudder bias system, (2) hydraulic power assist, (3) geared tab and (4) split rudder.
Depending on the design or regulatory requirements applicable to a given airplane design, reversible flight control systems may have single or dual mechanisms. The example shown in Figure 1 has a single flight control mechanism (i.e, single flight control run). Smaller general aviation airplanes designed to 14CFR Part 23 will have single flight control runs for each axis (pitch, roll and yaw). Reversible flight control systems in larger airplanes, such as corporate business jets, designed to 14CFR Part 25 will have dual flight control runs or have a separate backup system. This is due to the more stringent failure requirements required by regulation 25.671(c) in 14CFR Part 25. Dual flight control runs in each axis (roll, pitch and yaw) provide protection a jam or disconnection failure in any single run.
Design considerations for reversible flight control systems can be found in Flight Control System – Design Considerations. A discussion of design requirements can be found in Flight Control Systems – Design Requirements.